1. Technical Field
This invention relates to turbine engine rotor assemblies in general, and to blade outer tip seal apparatus in particular.
2. Background Information
A typical gas turbine engine includes a fan, compressor, combustor, and turbine disposed along a common longitudinal axis. The fan and compressor sections work the air drawn into the engine, increasing the pressure and temperature of the air. Fuel is added to the worked air and burned within the combustor. The temperature of the core gas flow increases as a result of the combustion. The magnitude of the increase depends upon several factors, including the amount of fuel added within the combustors. The combustion products and any unburned air, hereinafter referred to as core gas flow, subsequently powers the turbine and exits the engine producing thrust.
In most cases, the turbine comprises several stages each having a rotor assembly and at least one stationary vane assembly. The core gas flow causes the rotor assemblies to rotate, thereby enabling the rotor assemblies to do work elsewhere in the engine. The stationary vane assemblies located forward and/or aft of the rotor assemblies guide the core gas flow entering and/or exiting the rotor assemblies.
Under steady-state conditions at a given altitude, the power setting of an engine correlates to a particular flow rate of fuel injected into the combustors. The level Of thrust produced for the volume of fuel burned may be referred to as the "thrust specific fuel consumption" of the engine at that power setting. During transient periods, on the other hand, when the engine is accelerating from a first steady-state power setting to a second steady-state power setting, additional fuel is required to maintain the same level of thrust. Thus, the thrust specific fuel consumption of the engine decreases as well as the efficiency of the engine.
A significant cause of decreased engine efficiency is dissimilar thermal growth within the engine, for example between the blade tips of the turbine rotor assemblies and the shroud surrounding them. Core gas flow allowed to pass between the tips of the rotor assembly blades and the shroud does not cause the rotor assemblies to rotate and therefore does not add to the work done within the turbine. This undesirable extra clearance is most pronounced during rapid acceleration. Hard deaccelerations, on the other hand, can cause the shroud to contract more rapidly than the rotor assembly and potentially create an interference between the rotor blades and the shroud. Hence, there is a tension between minimizing the clearance between the blade tips and the shroud for performance sake, and maintaining clearance adequate to accommodate thermal expansion and contraction of the rotor assembly and shroud.
Performance aircraft powered by gas turbine engines require engines that are quick to respond to changes in the power setting of the engine. Power setting changes are typically accomplished by changing the fuel flow rate within the engine, although the method for determining maximum permissible changes varies. In some control schemes, the power produced by a turbine engine is limited by the core gas flow temperature within the turbine of that engine. Core gas flow temperature is used as a limiting factor to avoid exposing the turbine components to a temperature which would decrease the usable life of components below an acceptable level. Fuel flow rate, and therefore thrust produced, under a turbine temperature control scheme can be increased until maximum turbine temperature is reached.
A disadvantage of this approach is that the maximum possible thrust may not be available during a transient stage. During a hard acceleration, for example, the fuel flow may be increased dramatically until the maximum allowable turbine temperature is reached. The amount of time it takes to reach maximum allowable temperature is less, however, than the amount of time it takes the rotor assembly thermal growth to "catch up" to shroud thermal growth. As a result, the clearance between shroud and the rotor blade tips increases. The resulting inefficiency will decrease the available thrust until the engine reaches steady-state at which time the maximum available thrust will be produced. The time period between the call for maximum power and the time after which maximum power is available represents a lag in performance. A person of skill in the art will recognize that any lag in maximum power available is a serious detriment for a performance aircraft.
To avoid that undesirable lag, other control schemes use the pressure difference across the air inlet and the turbine exhaust to limit maximum permissible changes in power setting. Increasing the fuel flow rate within the combustors will cause an almost immediate increase in pressure within the turbine exhaust, and therefore in the pressure difference as well. As a result, the maximum possible thrust is available almost immediately. A disadvantage of this approach is that increasing the fuel flow rate to reach the desired pressure difference associated with maximum power during a transient period increases the core gas flow temperature within the turbine beyond the maximum permissible temperature associated with the desirable usable life of the turbine components. The extent to which the actual temperature exceeds the maximum permissible turbine temperature, and the duration of the exposure, depends upon the speed at which the turbine changes from transient state to steady state. Hence, both the performance and the life of the turbine components depend upon the thermal growth characteristics of the turbine rotor assemblies and shroud.